C.P. No. 1346
PROCUREMENT EXECUTIVE, MINISTRY OF DEFENCE
AERONAUTICAL RESEARCH COUNCIL
CURRENT PAPERS
Measurement of the Internal
Performance of a Rectangular Air-Intake
at a Mach Number of 0.9
C. 5. Brown and E. 1. Goldsmith
Aerodynamics Dept., R.A.E., Bedford
LONDON: HER MAJESTY’S STATIONERY OFFICE
1976
PRICE f/-80 NET
UDC 533.697.24 : 532.542.1 : 533.6.011.35
*CP No.1346 January, 1973
MEASUREMENT OF THE INTERNAL PERFORMANCE OF A RECTANGULAR AIR-INTAKE AT A MACH NUMBER OF 0.9
bY
c. s. Brown E. L. Goldsmith
SLJMMARY
A rectangular intake having variable geometry compression surfaces and
designed for supersonic operation has been tested at a Mach number of 0.9 and a Reynolds number based on intake entry height of approximately 106. Tests have been made on the intake in isolation and on the side of a fuselage.
Maximum mass flow is less than estimates based on geometric throat size.
The deficit corresponds to an effective reduction in throat area of about 3 per cent. Critical point pressure recovery is lower than predicted, but the diff- erence can be related to a 'turning loss' factor as was the case at sup&sonic speeds.
With the compression surfaces of the intake horizontal, performance was unaffected by angle of incidence within the range 0' to 10'. With the compres-
sion surfaces vertical however, there was some loss in maximum mass flow and pressure recovery at critical flow conditions at angles of incidence above about 4O. This was only partly alleviated by removal of the swept endwalls.
The presence of this particular fuselage imposed no obvious effect on the performance of the intake.
__--. - -.._ -~~ * Replaces RAE Technical Report 73020 - ARC 34702.
2
CONTENTS
I INTRODUCTION
2 DESCRIPTION OF THE TEST RIG
3 DETAILS OF THE MODEL
4 DETAILS OF THE TESTS 4.1 Test conditions
4.2 Instrumentation and data reduction
4.3 Accuracy
5 DISCUSSION OF RESULTS
5.1 Isolated intake
5.2 Intake mounted on a fuselage with compression surfaces horizontal
5.3 Intake mounted on the fuselage with compression surfaces vertical
9
5.4 Comparison between vertical intake and horizontal intake
5.5 Engine face distortion 6 CONCLUSIONS Notation References Illustrations Detachable abstract cards
9
9
10
11
12 Figures 1-25
Page
3
3
3
4
4 4 6 6 6
8
.
-
3
1 INTRODUCTION
An extensive programme of wind-tunnel tests has been carried out at
RAE, Bedford to investigate the internal performance of a particular rectangular
intake both in a uniform flow field and on a fuselage.
Details of the supersonic performance at Mach numbers between 1.6 and 2.5
are contained in Refs.1, 2 and 3. This Report presents the results from tests
made on the same intake at a Mach number of 0.9. Tests have been made on the
intake in isolation and also when installed on the side of a fuselage forebody
with the compression surfaces of the intake both horizontal and vertical. In
the case of the vertical intake the effect of removing the swept endwalls was
investigated.
2 DESCRIPTION OF THE TEST RIG
Figs.1 and 2 show the intake, duct and forebody assembled on the General
Intake Test Rig used in the 3ft wind tunnel at RAE, Bedford. This rig has been
described in Ref.4. It consists of a sting support, a calibrated mass flow
control and measuring unit, a hydraulic actuator system for moving the intake
compression surface ramps and an instrumented duct with interchangeable exit
plugs for controlling and measuring the intake bleed flow.
3 DETAILS OF THE MODEL
The intake is designed for supersonic operation and the model is that used
in Refs.1, 2 and 3. The ratio of height to width at the entry plane is 1.54 and
the geometry of the compression surface ramps, cowl and bleed are as shown in
Fig.3. The first compression surface has a fixed angle 61 of 10' and the shock
from the leading edge theoretically falls on the cowl lip at a free stream Mach
number of 2.43. The second compression surface is movable and is linked to the
rear ramp. In the configuration in which the intake is mounted on the fuselage
with its leading edge vertical, the two movable surfaces are connected to the
hydraulic actuator system on the intake test rig. However, when the intake is
assembled on the fuselage so that its leading edge is horizontal this arrangement
is not possible and the movable surfaces although still linked together in the
manner shown in Fig.3, are controlled by means of a manually operated lead screw.
The gap between the second and rear ramps forms a slot for bleeding the
boundary layer from the compression surfaces and this slot extends the whole
width of the intake. The bleed air is discharged through a duct into the free
stream. Bleed exit area is varied by means of interchangeable plugs. In
addition to zero bleed, tests were done at constant ratios of bleed exit area to
intake entry area of 0.05, 0.101 and 0.174.
4
Two different shapes of endwall were used in these tests. One was a full
swept endwall in which the leading edge coincides with the line joining the
leading edge of the intake to the cowl lip. The other is a minimal endwall
which is sufficient to contain the space under the second ramp at maximum 6 2'
but otherwise has a leading edge which is vertical at the cowl lip. Details of
the two shapes are shown in Fig.4.
The area distribution through the intake and duct for various values of
&2 is shown in Fig.5a while Fig.5b shows the area distribution through the
intake alone between the cowl lip and the rear hinge for those values of 62
used in the tests. The ratio of engine face cross-sectional area to maximum
intake capture area is 0.88 and the distance from the cowl lip to the engine
face is 9.89 times the intake height.
The nose and canopy only of the fuselage are represented. Fig.6 shows
details of the forebody including the relationship between the fuselage datum,
the intake datum and the nose cone centre line. The model is mounted in the
tunnel on the intake datum line so that c1 is the angle of incidence of the
intake relative to the wind-tunnel free stream.
4 DETAILS OF THE TESTS
4.1 Test conditions
The tests were all made in the 3ft wind tunnel at RAE, Bedford, utilising
the working section with top and bottom slotted walls. Tunnel conditions were
such as would give a free stream Mach number of 0.90 in the empty tunnel. The
effect on free stream Mach number of the intake test rig which has the rather
high blockage ratio of about 3 per cent is not known and no corrections for
wall constraint have been applied. The Reynolds number based on entry height
was approximately 106.
4.2 Instrumentation and data reduction
The standard mass-flow control and measuring unit used in these tests is
fitted with a cruciform rake having a total of 24 pitot tubes for measuring
total pressure at the engine face station. The tubes are disposed for area-
weighted averaging and the rake is rotatable to enable pressure surveys to be
made in greater detail'. Static pressure at the engine face is measured by using
four holes equally spaced around the circumference. Details of the character-
istics and calibration of this type of airflow meter are to be found in Ref.5.
The bleed duct contains twelve pitot tubes arranged in three rakes of four
tubes for measuring total pressure and three holes for measuring static pressure.
5
Pressure recovery is defined as:-
n pf pB 1 PO3
or c=nP, c P.
J (1)
1
where P co is the free stream total pressure and P. J
is the pitot pressure at
the jth tube in the rake at the engine face station in the case of Pf or in
the rake in the bleed duct in the case of PB; and n is in either case the
number of pressure points in the survey.
The ratio of engine face mass flow to maximum capture flow was computed
from:-
A a3 ( ) A, f
n
c KfAf 1 = nA M
e 00 (2)
f 1 + 0.2M2 5
co
where M. = J
local Mach number calculated from P. and pf; pf being the J
average static pressure at the engine face
Ae = maximum capture area
Af = area at the engine face
Kf = discharge coefficient obtained by calibration5.
Similarly bleed mass flow ratio was computed from:-
Am ( ) Ae B
= (3)
where P. J
now refers to pressures in the bleed duct and M. is the local Mach J
number calculated from P. J and pB9 the average static pressure in the bleed
duct. pb is the area of the bleed duct, in this case used without a discharge
coefficient.
Total mass flow ratio
(?jT = (tgf +(t$ l
(4)
The parameter used to define flow distortion at the engine face is DCbO,
defined as
DC6O = p60 - Pf)
9f (5)
where p60 is the mean total pressure in the worst 60' sector and qf is the
mean dynamic pressure at the engine face, derived from P f and pf'
4.3 Accuracy
Errors in the direct measurement of engine face and bleed duct total
pressure and therefore in pressure recoveries based on free stream total pressure
are thought to be small, not more than 0.1 per cent. Uncertainty associated with
the mean tunnel Mach number and more particularly with the calibration of the
airflow meter probably means that the error in mass flow ratio could be as much
as half a per cent.
l
5 DISCUSSION OF RESULTS
5.1 Isolated intake
Figs.7, 8 and 9 show the internal performance of the isolated intake at
three different values of bleed exit area, including the case of zero bleed, and
for a range of values of second ramp angle 62. Pitching the intake with its
compression surfaces horizontal had no effect on the pressure recovery - mass
flow characteristics within the range of incidence investigated.
Fig.10 summarises the variation with bleed exit area of maximum total and
maximum engine face mass flow, maximum pressure recovery and critical point
pressure recovery. The effect on performance of internal contraction in the
duct and of the location of the throat in the duct is well illustrated. At
values of second ramp angle 6 2 above about 5' where the minimum duct area is
in the region of the bleed slot, the bleed becomes inoperative at maximum mass
flow owing to the low internal pressure.
7
At values of 62 between 2O and So it is possible to obtain small amounts
of bleed flow but as bleed exit area is increased the intake soon chokes in the
region of the bleed slot. At values of A2 below 2' the minimum duct area is
at the rear hinge and bleed mass flow increases quite normally with increasing
bleed exit area. For all values of g2 maximum engine face mass flow is
practically independent of bleed exit area. For values of 62 above about 2'
the large bleed exit area (AB/Ae) = 0.10 results in a slight reduction in
maximum mass flow. This is not clearly understood.
The apparent minimum duct area has been derived from measured maximum mass
flow and is shown in Fig.:l, together with the geometric minimum duct area and
the inlet entry area at the cowl lip. The ratio of apparent to geometric minimum
duct area is approximately 0.97 except when the throat is at the rear hinge, in
which case the ratio falls to a minimum of 0.955.
Critical point pressure recovery shows a very slight increase with
increasing bleed exit area (Fig.10). This occurs with or without bleed flow
except at &2
equal to 0' where with a large bleed flow pressure recovery falls
off by about 2 per cent. Estimates of skin friction and interference losses
based on the work of Ref.6 indicate that at critical flow conditions there are
large additional losses. In reports of tests at supersonic speeds 192 these
additional losses have been ascribed to the turning of the flow in the duct
downstream of the cowl lip. In Fig.12 the losses in the present case are
plotted against w, the angle of turning between the front and rear movable
ramps. The correlation is again quite good, although the loss at 62 equal
to go seems to be rather high. Some explanation of this may be contained in
Fig.14. This shows the pressure distribution on the vertical centre line of the
duct obtained from a pitot rake near the rear hinge position as shown in Fig.3.
For each value of &2 and bleed exit area the pressure distribution is shown
for successive points on the pressure recovery-mass flow characteristic, pro-
gressing from a supercritical condition, through critical, to a subcritical
condition. At values of 62 of 5 0 and above, where the minimum duct area is
in the region of the bleed slot, the pressure distributions indicate that at
supercritical flow conditions the flow at the pitot rake is supersonic. Fig.13
shows a comparison of the pressure recovery indicated by the pitot tubes in the
middle of the duct when the intake is just supercritical, with the pressure
recovery calculated from the assumption that the Mach number at the rake is
derived from the area ratio At'Pk and Mt = 1.0.
8
The correspondence is quite good. This expansion over the rear ramp will
inevitably be followed by a normal shock system which in general will be prac- tically extinguished at the critical point. Most of the additional losses will be associated with separation occurring on the surface of the rear ramp. How- ever, at the relatively high 62 of go, followed by a rapidly expanding duct
downstream of the throat, there is probably still some significant shock loss at the critical flow condition.
5.2 Intake mounted on a fuselage with compression surfaces horizontal
Pressure recovery-mass flow characteristics for the intake when installed on the fuselage with its leading edge horizontal are shown in Fig.15. The data were obtained at a constant bleed unit area ($/A,) equal to 0.10 and for values of second ramp angle 62 equal to -loo, 0' and 5'. As was the case with the isolated intake varying the angle of incidence had no effect on these characteristics within the range investigated.
In Figs.lGa and 16b maximum total mass flow and maximum engine face mass flow are shown plotted against second wedge angle 62 for the intake alone and
for the intake on the fuselage. Estimates of mass flow based on intake internal geometry have been calculated and are included for comparison. The presence of the fuselage causes a reduction in maximum total mass flow of about 1.5 per cent.
The difference in approach Mach number between the isolated intake and the intake on the fuselage is not known; all mass flows have been referred to nominal free stream Mach number. However, the difference is hardly likely to be the cause of this reduction in total flow because there is no corresponding reduction in engine face mass face. The deficit is entirely in the bleed flow quantity and is more probably caused by a change in bleed exit conditions due to the presence of the fuselage.
Measured maximum engine face mass flow falls below the quantity estimated from the intake internal geometry because of the apparent throat size referred to in section 5.1. The difference between the measured maximum total mass flow and the estimate based on the inlet area at the cowl lip has one of two origins depending on the particular value of 62. For values of 62 above about 2' it is due largely to choking at the bleed inlet and consequent failure of the bleed to operate. At lower values of 62 it is due simply to insufficient bleed exit area.
The negligible effect of the fuselage flow field on intake pressure recovery is illustrated in Figs.16c and 16d which show the variation with d2
9
of both critical point and maximum pressure recovery for the isolated intake and
the intake on the fuselage. Estimates of skin friction and interference losses
based on Ref.6 are also shown on this figure to illustrate the large additional
pressure recovery losses referred to in section 5.1.
5.3 Intake mounted on the fuselage with compression surfaces vertical
Figs.17 to 22 show pressure recovery-mass flow characteristics for the intake when mounted on the fuselage with its leading edge vertical. The intake was fitted with either two swept endwalls or two unswept endwalls. Details of
the endwall shapes are shown in Fig.4. For each endwail configuration data
are presented for a constant second ramp angle equal to -loo, with bleed exit
areas MB/A, 1 of 0.05, 0.10 and 0.174, and a range of incidence of the intake-
fuselage assembly from zero to 12'.
In Fig.23 the variation of mass flow and pressure recovery with bleed exit
area is summarised for both endwall configurations. Removing the swept end- walls reduces the large loss in maximum pressure recovery which occurs at an
incidence of 12'.
5.4 Comparison between vertical intake and horizontal intake
The mass flows and pressure recoveries of the various configurations of
the intake when mounted on the fuselage are shown in Fig.24 for a second ramp angle of -10' and a bleed exit area (AB/Ae) equal to 0.10. The indications are that at low angles of incidence (01 = Co to 4') there is little difference between any of the configurations. However, the intake with its compression surfaces vertical is affected by incidences above 4'. In particular the configuration with two swept endwalls suffers a significant loss in maximum pressure recovery at an incidence of 12'.
5.5 Engine face distortion
Distortion at the engine face expressed as DC 60
is sho-m in Fig.25. The maximum subcritical value of DC 60 is also shown, This is rarely greater than the value at the critical point and in any case always occurs within 5 per cent
of maximum mass flow. At low angles of incidence, up to about 4', the value of
DC60 is in the region -0.25 to -0.35 for all configurations. However, at angles of incidence beyond 4' the intake with the vertical leading edge shows some deterioration in flow uniformity whereas with the Leading edge horizontal the uniformity is apparently unaffected.
10
6 CONCLUSIONS
Measurements have been made of the internal performance of a rectangular
intake having variable geometry and designed for supersonic operation. The
measurements were made at a Mach number of 0.9 and a Reynolds number based on
intake entry height of approximately 106. The intake has been tested through a
range of angle of incidence of 0' to 10' in isolation in a uniform flow field
and in the environment generated by a particular shape of fuselage.
Maximum mass flow is slightly less than would be estimated on the basis
of geometric throat size; the deficit corresponds to a reduction in throat size
of about 3 per cent.
Internal contraction, where it is located in the region of the bleed slot,
leads to a failure of bleed operation at maximum mass flow.
Critical point pressure recovery is well below that predicted after taking
into account duct losses. However, as was the case at supersonic speeds, these
additional losses correlate well with the turning angle between the front and
rear movable ramps.
With the compression surfaces of the intake horizontal there was no
deterioration in performance with increase in angle of incidence; nor was the
performance impaired by the presence of this particular fuselage.
With compression surfaces vertical, the intake on the fuselage, suffered
some loss in mass flow and pressure recovery at angles of incidence above 4 0
This was only partly alleviated by removing the swept endwalls.
11
% % AR Af 4r K
L M
P
pR q P X a
% &2 w
NOTATION
bleed exit area bleed duct cross-sectional area
cross-sectional area at the rake station near the rear hinge
engine face cross-sectional area cross-sectional area at station X
discharge coefficient distance from the cowl lip to the rear hinge Mach number
total pressure pitot pressure at the rake station near the rear hinge
dynamic pressure static pressure distance downstream of cowl lip angle of incidence of intake relative to free stream
angle between first compression surface and free stream ahead of intake angle between first and second compression surfaces
angle between front and rear movable ramps
Subscripts
C at critical flow conditions
B in the bleed e in the intake entry plane f at the engine face i at the cowl lip t at the section of duct minimum area T total, i.e. engine plus bleed
X at station X al in the free stream
12
No. Author -
1 C.S. Brown
E.L. Goldsmith
2 C.S. Brown
E.L. Goldsmith
3 C.S. Brown
E.L. Goldsmith
4 E.L. Goldsmith
5 I. McGregor
6 J. Seddon
REFERENCES
Title, etc.
Measurement of the internal performance of a rectangular
air intake with variable geometry. Part I.
ARC CP No.1243 (1971)
Idem. Part II. The effect of incidence.
ARC CP No.1292 (1971)
Measurement of the internal performance of a rectangular
air intake mounted on a fuselage at Mach numbers from
1.6 to 2.
ARC CP No.1291 (1972)
Variable geometry intakes at supersonic speeds. Some
techniques and some test results.
AGARD CP 34 (1968)
The characteristics and calibration of two types of
airflow metering device for investigating the performance
of model air intakes.
RAE Technical Report 71212 (ARC 33894) (1971)
Boundary layer interaction effects in intakes with
particular reference to those designed for dual subsonic
and supersonic performance.
ARC R & M 3565 (19663
\
Cod internal angle 1J”
/
Cowl external angle 2-149 in
b-689 in lb9*9mm
Fig.3 Geometry of compression surface ramps, cowl and bleed
9-017 in
102 mm
c
Lea&q edge
Swept endwafl Unswept endwall
E E cu . s
a 66-lmm
e
Leading edqe
Fig.4 Details of endwalls
2.2
2-o
1.6
l-6 A, Ai I.4
I I I I I I I
- - - . a-
0-I 03 o-4 o-5
X (enqme face) .
Fig.Sa Area distribution through intake
I I
o-5 x I-0 0
Fig. 5b Area distribution through intake between cowl lip and rear ramp hinge
I-0
o-95
Pf
P,
0.90
O-85
I-0
PfJ o-9 p, O-8
o-7
O-6
O-7
i I I I I
03 o-4 (
0-j o-4 C
.
I-O
o-95
Pf
P,
o-90
0.65 1
,
*a3 0
0.04
ii- t B o-03
O-02
0.01 0
1
Fig.8 Variation of pressure recovery with mass flow. Isolated in take
I.0
Pf
PO0 o-95
pa O-8 - PO0 0.7
r ___I___ ___^.___ .- .__-. - --.--._-- -1
O-6
o-5
06 o”- IO0 AB (> -b,
Fig.9 Variation
o-95
0*9c
0.08
0.06 .
of pressure re Isolated intake
. ,
045
bo 0 A,
6, 0’ 625” I
A, 0-I ( > Ae
O-90 1 I I 1 , .
Fig.10 Variation of mass flow and pressure recovery bleed exit area. Intake alone. cL=O*-IO*
0.75
O-70
0.65
O-60 At A,
053
O-50
o-45
O-40
Ceometrlc ------
Derived from measured mass flow:-
0 0.05 0.10 0 X A
I I I 1
0 2 4 6 8 6,” 10
Fig. II hct minimum area derived from measured mass flaw.
Isolated intake
. * h
0
./
p”
0’
/ I
/ .
/ I I I I
- - - I - . -a
0 o-2 o-4 O-6 0.6
Anqle between front and rear movable ramPs O(radians)
. \ I
O-96
o-94
0.92
o-90
Calculated assuminq sonic throat
Measured o
I I I I 2 4 6 8
Second ramp anqle 6;
Fig.12 Correlation between flow turning Fig. 13 Pitot pressure at rear hinge loss and angle between front and rear position. Critical flow conditions. movable ramps. Isolated intake. Zero bleed Isolated intake
61 9”
6, 7”
62 5”
bt 2’
62 0’
I I
1 0 K
A
Cod side
Fig.14 Pressure distribution on duct vertical centre line at rem hinge position
b L
L n
I- 0
0 g Q)
o-95
O-90
00
0
0.6 4 P,
o-4
O-2
1 I , 1
0.5 0.4
I.0
Pf
0 a,
0.95
Fig.15 Variation of pressure recovery with. mass Intake with swept endwalls on fuselage with
O-70
0.65 ho
0 A eT max
0~60
\ \
\ \
I \
\
\ \ \
\ \
\ I
\ \ \ \ \
\
\ \
\ \
7
a
+kin friction
O-65
O*bO
o-55
ox
Istimated from ntakc qeometry--- leasured :- ntake alone Q ,ntake on fuselage A
0 5 62’ IO b
rice loss Ref 6
I
5 6 d
D
Fig. 16 a-d Variation of pressure recovery and maximum mass flow with 6,. Intake alone II on fuselage with leading
edge horizontal (*‘I&) = o- IOI
L ” r
I4
o-95
o-9(1
0.85 I
I-0
o-95 P f PO0
0%
O-85 ii 0.j 0.4 O-6 O-7
0.05
0.3 o-4 O-5 AS O-6
0
O-7 A
ef
Fig. I7 variation gf pressure recovery with moss flow. Intake with swept endwalls on fuselage with LE vertical
I-0
o-9: Pf
0*9(
O-85
0.9
pB O-8 P,
o-7
0-b
o-5 I I I I
3 o-4 0.7
I-0
o-95
Pf
Ki
o-90
0.8518 0-J 0.4 0.7
O-05, 0.04
o*o’J
0.02
o-01 n
6, = - IO” A*
1 0 A, = 0.05 a zoos 4” 0 &=8” x cx=12” A
1 I “0-j 0.4
I I I O-6 0.7
Fig. 18 Variation of pressure recovery with mass flow. Intake with unswept endwalls on fuselage with L E vertical
. .
I-0
o-95 P t L
o-90
O-85
I-0
o-9
P, O-6 PO0 O-7
,- O-6
o-5 1 I I I 0-j 0.4 O-5 A, O-6 O-7
0 A Q T
v-
, -
I-O
o-95
Pf
P,
O-9(
O-8!
O-10
o-09
0.06
o-05
Fig. I9 Variation of pressure recovery with Intake with swept endwalls on fuselage with
o-95 !i
O-85 I
0.6
0.5
62 = - IO0
I I
3 o-4 0-5 A, O-6 O-7
0 A, 1
cx= 0" and 4' 0
I ot = 8" X
oc = 12” A
I I 5 o-4 o-5
Cl B ii-6 d-7 A, t
O-95 3
iii
O-PO
O-85 I
Fig.20 Variation of pressura recovery with mass flow. Intake with unswept endwoils on fuselage with C E vertical
. .
I-0
o-95 P f pop
o-90
0.85
74 A lb
I I I I I
09 O-4 A, O-5 ( \
O-6 0.7 7
o-9
P, O-8 p, o-7
O-6
@ I.0
0.95
Pf
L
0.90
0.85 I I I I
03 O-6 O-7
O-08 1 I I I I
03 O-5 O-6 O-7
Fig.2 I Variation of pressure recovery with mass flow. intake with swept bdwalls on fuselage with LE vertical
I-0
0.95
!k PCQ
o-90
O-85
I-0
o-95 P f pap
O-90
O-85 i I I 1 ,
03 Q-4 O-5 A, O-6 LA
O-7
62 = -10” *f!
0 A, = o-174
o-5 O-4
O-6
0.5 I 1 I I I 0.5 0.4 O-5 0 A, 0-b 0-7
A, T
O-001 I 1 I I 0-J O-4 O-5 A, O-6
0
O-7
AC f
Fig.22 Variation of pressure recovery with moss flow. Intake with unswept endwalls on fuselage with LE vertical
O-65
0.60
d o’and 4”
4 A 4
/- ---+ / +,L-x
h--A-.-.-d
I
01 2
02 e
i 01% 0.2
X
+
Ae -0 %?!2 Wept
A e I Unswept Z
AQ Swept endualls -
Swept x Unwept +
oc 12”
X
/
+ +- --+-RR
x--x-, I ’
w(
-’ max
1
Fig.23 Variation of mass flow and pressure recovery with bleed exit area.
Intake on fuselage. LE vertical. &=- IO0
0.65
O-60
LE vertlcal wept endwalls LE vertwl unwept enduralls ----- LL horizontal wept endulalls -----
*a 6 e t max
*oo
0 A ef max
i
Pf 0 ii max
Pf
0 L C
Fig.24 Comparison of mass flow and pressure recovery. Intake on fuselage with LE horizontal and
vertical. 6, = --IO” (2) =O*IO
t
- 0.5
- 0.4
- 03
- o-2
- 0-I
0
- 0.4
- o-3
- 0.2
- 0-I
0
0 4 OL” 8 I2 0 4 ado 8 12
I I I
0 4 oc” 8 I
Intake on fuselaqe L I vertical
Wept endwalls 6, - IO”
-- 4 d” 0 I
Intake on fuselage L E vertical
tInwet enhualls 2-
e 0 0
Y Y Y Y
+
+ + +
X X X 1
1 1
4 oc 8 I2 A
0 $ O*lOi
Y Y Y Y :
X
+ : r
I I
4 ac” 8 I2
Intake alone 6, 0" Intake on fuselaqe (L E horizontal)
Fig.25 Engine face distortion coefficient .
ARC CP No. 1346 January 1973
533.697.24 : 532542.1 . 533.6 011.35
Brown, C. S. Goldsmith, E. L.
MEASUREMENT OF THE INTERNAL PERFORMANCE OF A RECTANGULAR AIR-INTAKE AT A MACH NUMBER OF 0.9
A rectangular mtake having variable geometry compression surfaces and destgned for super- sonic operation has been tested at a Mach number of 0 9 and a Reynolds number based on mtake entry height of approximately 10‘. Tests have been made on the mtake in isolation and on the side of a fuselage.
Maximum mass flow is less than e&mates based on geometrrc throat sire. The deficit corresponds to an effective reductton m throat area of about 3 per cent Cnticai pomt pressure recovery is lower than predicted, but the dtfference can be related to a ‘turnmg loss’ factor as was the case at supersonic speeds.
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ARC CP No. 1346 January 1973
533.697 24 . 532.542.1 : 533.6.011.35
Brown, C. S. Goldsmtth, E. L.
MEASUREMENT OF THE INTERNAL PERFORMANCE OF A RECTANGULAR AIR-INTAKE AT A MACH NUMBER OF 0.9
A rectangular intake havmg vartable geometry compression surfaces and designed for super- sonic operation has been tested at a Mach number of 0.9 and a Reynolds number based on mtake entry height of approximately 10’. Tests have been made on the intake m isolation and on the side of a fuselage.
Maxlmum mass flow is less than estimates based on geometric throat sire. The deficit corresponds to an effective reduction in throat area of about 3 per cent. Cnticai pomt pressure recovery IS lower than predrcted, but the difference can be related to a ‘tummg loss’ factor as was the case at supersonic speeds.
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ARC CP No. 1346 January 1973
533.691.24 : 532.542 1. 533.6 011.35
Brown, C S Goldsmith, E L.
MEASUREMENT OF THE INTERNAL PERFORMANCE OF A RECTANGULAR AIR-INTAKE AT A MACH NUMBER OF 0.9
A rectangular mtake havmg variable geometry compression surfaces and designed for super- some operatton has been tested at a Mach number of 0.9 and a Reynolds number based on intake entry height of approximately 106. Tests have been made on the intake in isolation and on the side of a fuselage
Maxmmm mass flow is less than esttmates based on geometrtc throat sire. The deficit corresponds to an effective reduction in throat area of about 3 per cent Crittcal point pressure recovery 1s lower than predtcted, but the difference can be related to a ‘turnmg loss’ factor as was the case at supersonic speeds.
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With the compressron angle of incidence wrthm however, there was some flow conditrons at angles removal of the swept endwalls.
The presence of thus partrcular the mtake.
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With the compressron surfaces of the mtake horizontal, performance was unaffected by angle of mcidence within the range 0” to 10” With the compressron surfaces vertical however, there was some loss in maximum mass flow and pressure recovery at crrtical flow condrtions at angles of incidence above about 4O removal of the swept endwalls.
This was only partly alleviated by
I With the compression surfaces t angle of mcrdence within
t however, there was some flow conditions at angles removal of the swept endwalls.
The presence of this particular fuselage imposed no obvrous effect on the performance of the intake. I The presence of this partrcular
the intake.
C.P. No. I346
Q Crown cepyright
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C.P. No. 1346 ISBN 611 470981 5